CN110321598A - A kind of Spacecraft Relative Motion Analytical Solution method under J2 perturbation conditions - Google Patents
A kind of Spacecraft Relative Motion Analytical Solution method under J2 perturbation conditions Download PDFInfo
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Abstract
The invention discloses a kind of Spacecraft Relative Motion Analytical Solution methods under J2 perturbation conditions, the space orbit evolution task that relative motion can be carried out for the Spacecraft formation under the conditions of the aspherical J2 Gravitational perturbation of the earth/cluster can be achieved with Spacecraft formation/cluster relative track movement state parsing evolution by constructing relative motion state-transition matrix.Its evolution that absolute orbit is carried out with Spacecraft formation/cluster member J2 perturbation absolute orbit kinetics equation, by to the absolute orbit difference between member and having carried out coordinate transform and obtaining Differential Equations of Relative Motion using relative orbit state as variable, the differential equation is carried out by Taylor series expansion method to solve acquisition relative motion state-transition matrix and analytic solutions.
Description
Technical field
The invention belongs to Spacecraft formation dynamics of relative motion modeling technique fields, and in particular to a kind of J2 perturbation item
Spacecraft Relative Motion Analytical Solution method under part.
Background technique
With the development of space technology, modern space mission is increasingly intended to diversification, complicates, and passes through multiple space flight
The function and application that single spacecraft is difficult to realize is realized in device formation flight, is the important trend in aerospace engineering field.Boat
Its device formation flight realizes that the premise for completing particular task is to maintain certain geometric configuration, carries out the premise of Formation keeping control
It is to be modeled to formation dynamics of relative motion.
Currently, existing Spacecraft Relative Motion kinetic model is various, and various models are often subject between spacecraft
Use scope, orbital eccentricity, the type for modeling perturbed force and a variety of constraints and limitation that integrate required amount of calculation, and
The problem of can be only applied to feature.It is often difficult to directly judge the superiority and inferiority of some kinetic model, only in specific task field
See whether the model is applicable under scape.
Earliest dynamics of relative motion model by Clohessy and Wiltshire propose for establish space station or
The C-W equation in " Rendezvous " stage during track assembling Large-scale satellite.The model using the time as independent variable,
The relative position of two spacecrafts and the variation of speed are described, suitable for the opposite fortune of near-circular orbit in the case of being closer
It is dynamic.Then, it is linearized for the nonlinear problem of the relative motion on arbitrary ellipse track, Tschauner and Hempel,
It has derived using true anomaly as the inearized model of independent variable and its state-transition matrix.
CW equation and TH equation are the inearized models based on two-body problem, do not consider the effect of perturbative force,
Other scholars attempt that atmospheric drag perturbation, perturbation of earths gravitational field and life are added in kinetic model on this basis respectively
Three-body Gravitational perturbation etc..Since under Perturbation Effect, spacecraft absolute orbit element has slow alteration matter, some authors draw
Enter the state parameter of another description Spacecraft Relative Motion, referred to as relative light intensity (ROE), it is two spacecrafts six
The linearly or nonlinearly combination of orbital tracking.The model of Gim and Alfriend exploitation considers J2 perturbation and the rail of the compression of the Earth
Road eccentricity has obtained the state-transition matrix based on ROE.Rich Yin Rui etc. on this basis, considers the aspherical gravitation of the earth, big
The influence of the perturbative forces such as atmidometer and three-body gravitation is added to corresponding perturbing term for relative motion model.
Spacecraft Formation Flying orbits controlling and it is motor-driven in, relative light intensity description it is not intuitive enough, it is still desirable to
It is converted into real-time relative position and speed.Kechichian derives one group of nonlinear differential equation, describes pursuit spacecraft
Dynamics relative to passive space vehicle reference frame, it is contemplated that J2 perturbation shadows of highly elliptic orbit eccentricity and the earth
Ring, subsequent Threon etc. further simplifies equation, obtain numerical integration as a result, but related research result all do not obtain
Relative motion analytic solutions under J2 perturbation conditions.
Therefore, in the prior art, the dynamic analysis under J2 perturbation is carried out as state by relative light intensity to model
Orbital Evolution is carried out or by relative position and speed as the state progress J2 lower phases that perturb to Spacecraft formation relative motion
Numerical integration evolution is carried out to movement, lacks the item perturbation relative motion Analytical Solution using relative position and speed as the J2 of state
Method.
Summary of the invention
Above-mentioned the deficiencies in the prior art are directed to, the purpose of the present invention is to provide spacecrafts under a kind of J2 perturbation conditions
Relative motion Analytical Solution method can shift square by the state that this method obtains in the case where excessively not increasing calculation amount
Battle array, can be achieved with the evolution of Spacecraft formation relative motion track under J2 perturbation conditions.
In order to achieve the above objectives, The technical solution adopted by the invention is as follows:
Spacecraft Relative Motion Analytical Solution method under a kind of J2 perturbation conditions of the invention, comprises the following steps that
1) the absolute orbit kinetics equation comprising J2 perturbations of two spacecrafts is established under inertial coodinate system respectively；
2) difference is carried out to the absolute orbit kinetics equation under above-mentioned two spacecraft J2 perturbations, obtained under inertial system
Spacecraft taken the photograph dynamics of relative motion equation；
3) transformation matrix of coordinates of inertial coodinate system i to orbital coordinate system l (LVLH) is established；
4) dynamics of relative motion equation is taken the photograph to the spacecraft under inertial system using the transformation matrix of coordinates of foundation to carry out
Coordinate transform obtains being taken the photograph dynamics of relative motion equation under track system；
5) using under track system relative position and speed as quantity of state, moved relative motion is taken the photograph under established track system
Mechanical equation is organized into SYSTEM OF LINEAR VECTOR differential equation form；
6) the SYSTEM OF LINEAR VECTOR differential equation is solved, state-transition matrix is obtained.
Further, in the step 1) in the case where establishing inertial coodinate system under considering the aspherical J2 perturbation conditions of the earth two
The motion dynamics equations of spacecraft are as follows:
Wherein, rtAnd vtThe respectively relative position under 1 inertial system of spacecraft and speed, rcAnd vcRespectively spacecraft 2 is used
Relative position and speed under property system, g (rt) and g (rc) be respectively spacecraft 1 and 2 normal gravity:
Wherein, μ is Gravitational coefficient of the Earth, ReqIt is earth mean radius, n=[0,0,1]TIt is earth's axis under inertial system
The unit vector in direction,WithIt is r respectivelytAnd rcUnit vector, J2=0.00108263 is ground
The aspherical perturbation coefficient of ball, | | | | indicate the operation of modulus value.
Further, the spacecraft under inertial system is obtained in the step 2), and to be taken the photograph dynamics of relative motion equation as follows:
Wherein, r and v is respectively the relative position under inertial system between two spacecrafts and speed.
Further, the inertial coodinate system established in the step 3) to orbital coordinate system (LVLH) transformation matrix of coordinates such as
Under:
Wherein,Indicate transformation matrix of coordinates of the i system to l system, T representing matrix transposition operation.
Further, under the track system established in the step 4) to be taken the photograph dynamics of relative motion equation as follows:
Wherein, subscript l indicates the projection under orbital coordinate system l, Ω×It is the multiplication cross square of the orbit angular velocity ω of spacecraft 2
Battle array,Indicate g (r) in rcThe partial derivative at place；
Wherein,I3×3Indicate 3 × 3 unit matrix.
Further, in the step 5) using under track system relative position and speed as quantity of state, i.e. x (t)=[rl,vl]T,
By under track system by take the photograph dynamics of relative motion equation be organized into linear differential equation form it is as follows:
Wherein, F (t) is sytem matrix, and form is as follows:
Further, the step 6) specifically includes: formula (8) is solved according to lineary system theory, as a result as follows:
X (t)=Φ (t) x (0) (10)
Wherein, x (0)=[r0 l,v0 l]TIt is given initial relative position and speed, Φ (t) is the corresponding state of t moment
Transfer matrix, concrete form are as follows:
Φ (t)=et·F(t) (11)。
Wherein, e is natural constant.
Beneficial effects of the present invention:
The present invention can carry out opposite fortune for the Spacecraft formation under the conditions of the aspherical J2 Gravitational perturbation of the earth/cluster
It is opposite to can be achieved with Spacecraft formation/cluster by building relative motion state-transition matrix for dynamic space orbit evolution task
The parsing of track motion state is developed.It is carried out with Spacecraft formation/cluster member J2 perturbation absolute orbit kinetics equation
The evolution of absolute orbit, by the absolute orbit difference between member and carried out coordinate transform obtain with relative orbit state
As the Differential Equations of Relative Motion of variable, the differential equation is carried out by Taylor series expansion method to solve acquisition relative motion
State-transition matrix and analytic solutions.
Detailed description of the invention
Fig. 1 is the modeling method of the invention schematic diagram；
Fig. 2 a is X-direction relative position estimation error curve figure；
Fig. 2 b is Y direction relative position estimation error curve figure；
Fig. 2 c is Z-direction relative position estimation error curve figure；
Fig. 3 a is X-direction relative velocity estimation error curve figure；
Fig. 3 b is Y direction relative velocity estimation error curve figure；
Fig. 3 c is Z-direction relative velocity estimation error curve figure.
Specific embodiment
For the ease of the understanding of those skilled in the art, the present invention is made further below with reference to embodiment and attached drawing
Bright, the content that embodiment refers to not is limitation of the invention.
Spacecraft Relative Motion Analytical Solution method under a kind of J2 perturbation conditions of the present invention, for passing through in the prior art
Relative light intensity carries out the dynamic analysis modeling under J2 perturbation as state and carries out track to Spacecraft formation relative motion
Develop or relative motions progress numerical integration evolutions under J2 perturbations carried out as state by relative position and speed, lack with
Relative position and speed as the J2 item of state perturb relative motion Analytical Solution method the problem of.Method of the invention by pair
Absolute orbit difference between spacecraft has simultaneously carried out coordinate transform and obtains relative motion using relative orbit state as variable
The differential equation carries out the differential equation by Taylor series expansion method to solve acquisition relative motion state-transition matrix and parsing
Solution, the track that can be suitable for Spacecraft formation/cluster relative motion aerial mission, which parses, to be developed.
Referring to Fig.1, it is described as follows:
1, in the case where establishing inertial coodinate system under considering the aspherical J2 perturbation conditions of the earth two spacecrafts track power
It is as follows to learn equation:
Wherein, rtAnd vtThe respectively relative position under 1 inertial system of spacecraft and speed, rcAnd vcRespectively spacecraft 2 is used
Relative position and speed under property system, g (rt) and g (rc) be respectively spacecraft 1 and 2 normal gravity.
2, difference is carried out to the absolute orbit kinetics equation under two spacecraft J2 perturbations, obtains the boat under inertial system
Its device is taken the photograph dynamics of relative motion equation:
Wherein, r and v is respectively the relative position under inertial system between two spacecrafts and speed.
3, the inertial coodinate system established is as follows to orbital coordinate system (LVLH) transformation matrix of coordinates:
Wherein,Indicate transformation matrix of coordinates of the i system to l system, T representing matrix transposition operation.
4, under the track system established to be taken the photograph dynamics of relative motion equation as follows:
Wherein, subscript l indicates the projection under orbital coordinate system l, Ω×It is the multiplication cross square of the orbit angular velocity ω of spacecraft 2
Battle array,Indicate g (r) in rcThe partial derivative at place.
5, using under track system relative position and speed as quantity of state, i.e. x (t)=[rl,vl]T, by being taken the photograph under track system
The form that dynamics of relative motion equation is organized into linear differential equation is as follows:
Wherein, F (t) is sytem matrix.
6, formula (5) is solved according to lineary system theory, as a result as follows:
X (t)=Φ (t) x (0) (6)
Wherein, x (0)=[r0 l,v0 l]TIt is given initial relative position and speed, Φ (t) is the corresponding state of t moment
Transfer matrix, concrete form are as follows: Φ (t)=et·F(t), wherein e is natural constant.
The example of the method for the present invention: illustrate case verification of the invention in conjunction with Fig. 2 a to Fig. 3 c, set following design conditions
And technical parameter:
1) preliminary orbit parameter of the satellite A under inertial coodinate system are as follows:
[5023.5585km,5023.5585km,0km,-1.8109km/s,1.8109km/s,7.0411km/s]；
2) preliminary orbit parameter of the satellite B under inertial coodinate system are as follows:
[5023.4575km,5023.6791km,0.4700km,-1.8108km/s,1.8104km/s,7.0413km/s]；
3) simulation time 27000 seconds；
4) the state update cycle is 10 seconds；
Design conditions and technical parameter based on relative motion analytic method of the invention and above-mentioned setting use
Mathematics software carries out simulating, verifying.It is relative position and speed between two spacecrafts respectively as shown in Fig. 2 a to Fig. 3 c
Evolution error curve is spent, it is found that three axis relative position evolution errors are within 1m, relative velocity error exists You Tuzhong curve
Within 0.001m/s, relative orbit evolution precision is very high.
Therefore, using the method for the present invention, it can realize that J2 perturbation relative track movement parsings are built in cartesian coordinate system
Mould and solution realize that the precise orbit of Spacecraft Relative Motion develops.
There are many concrete application approach of the present invention, the above is only a preferred embodiment of the present invention, it is noted that for
For those skilled in the art, without departing from the principle of the present invention, it can also make several improvements, this
A little improve also should be regarded as protection scope of the present invention.
Claims (7)
1. a kind of Spacecraft Relative Motion Analytical Solution method under J2 perturbation conditions, which is characterized in that comprise the following steps that
1) the absolute orbit kinetics equation comprising J2 perturbations of two spacecrafts is established under inertial coodinate system respectively；
2) difference is carried out to the absolute orbit kinetics equation under above-mentioned two spacecraft J2 perturbations, obtains the boat under inertial system
Its device is taken the photograph dynamics of relative motion equation；
3) transformation matrix of coordinates of inertial coodinate system i to orbital coordinate system l is established；
4) dynamics of relative motion equation is taken the photograph to the spacecraft under inertial system using the transformation matrix of coordinates of foundation and carries out coordinate
Transformation, obtains being taken the photograph dynamics of relative motion equation under track system；
5) using under track system relative position and speed as quantity of state, dynamics of relative motion will be taken the photograph under established track system
Equation is organized into SYSTEM OF LINEAR VECTOR differential equation form；
6) the SYSTEM OF LINEAR VECTOR differential equation is solved, state-transition matrix is obtained.
2. Spacecraft Relative Motion Analytical Solution method under J2 perturbation conditions according to claim 1, which is characterized in that
In the case where establishing inertial coodinate system under considering the aspherical J2 perturbation conditions of the earth, the track of two spacecrafts is dynamic in the step 1)
Mechanical equation is as follows:
Wherein, rtAnd vtThe respectively relative position under 1 inertial system of spacecraft and speed, rcAnd vcRespectively 2 inertial system of spacecraft
Under relative position and speed, g (rt) and g (rc) be respectively spacecraft 1 and 2 normal gravity:
Wherein, μ is Gravitational coefficient of the Earth, ReqIt is earth mean radius, n=[0,0,1]TIt is earth rotation axis direction under inertial system
Unit vector,WithIt is r respectivelytAnd rcUnit vector, J2=0.00108263 is that the earth is non-
Spherical perturbation coefficient, | | | | indicate the operation of modulus value.
3. Spacecraft Relative Motion Analytical Solution method under J2 perturbation conditions according to claim 1, which is characterized in that
The spacecraft under inertial system is obtained in the step 2), and to be taken the photograph dynamics of relative motion equation as follows:
Wherein, r and v is respectively the relative position under inertial system between two spacecrafts and speed.
4. Spacecraft Relative Motion Analytical Solution method under J2 perturbation conditions according to claim 1, which is characterized in that
The inertial coodinate system established in the step 3) is as follows to orbital coordinate system transformation matrix of coordinates:
Wherein,Indicate transformation matrix of coordinates of the i system to l system, T representing matrix transposition operation.
5. Spacecraft Relative Motion Analytical Solution method under J2 perturbation conditions according to claim 1, which is characterized in that
Under the track system established in the step 4) to be taken the photograph dynamics of relative motion equation as follows:
Wherein, subscript l indicates the projection under orbital coordinate system l, Ω×It is the multiplication cross matrix of the orbit angular velocity ω of spacecraft 2,Indicate g (r) in rcThe partial derivative at place；
Wherein,I3×3Indicate 3 × 3 unit matrix.
6. Spacecraft Relative Motion Analytical Solution method under J2 perturbation conditions according to claim 1, which is characterized in that
In the step 5) using under track system relative position and speed as quantity of state, i.e. x (t)=[rl,vl]T, by under track system by
Take the photograph dynamics of relative motion equation be organized into linear differential equation form it is as follows:
Wherein, F (t) is sytem matrix, and form is as follows:
7. Spacecraft Relative Motion Analytical Solution method under J2 perturbation conditions according to claim 6, which is characterized in that
The step 6) specifically includes: formula (8) solved according to lineary system theory, as a result as follows:
X (t)=Φ (t) x (0) (10)
Wherein,It is given initial relative position and speed, Φ (t) is the corresponding state transfer square of t moment
Battle array, concrete form are as follows:
Φ (t)=et·F(t) (11)。
Wherein, e is natural constant.
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113282096A (en) * | 2021-06-04 | 2021-08-20 | 中国人民解放军战略支援部队航天工程大学 | Control method for relative position nonlinear error of geostationary orbit game spacecraft |
| CN113761809A (en) * | 2021-11-08 | 2021-12-07 | 南京航空航天大学 | Passive detection orbit determination method based on deep neural network |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106628257A (en) * | 2016-09-28 | 2017-05-10 | 西北工业大学 | Method for keeping relative motion orbit of near-earth spacecraft in earth's perturbation gravitational field |
| CN109582039A (en) * | 2019-01-09 | 2019-04-05 | 北京空间飞行器总体设计部 | A kind of optimal reconfiguration method under the J2 perturbation using Relative Navigation information |
| CN109613932A (en) * | 2019-01-09 | 2019-04-12 | 北京空间飞行器总体设计部 | Optimal continuous thrust control method under a kind of J2 perturbation using Relative Navigation information |
-
2019
- 2019-06-10 CN CN201910494778.5A patent/CN110321598B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106628257A (en) * | 2016-09-28 | 2017-05-10 | 西北工业大学 | Method for keeping relative motion orbit of near-earth spacecraft in earth's perturbation gravitational field |
| CN109582039A (en) * | 2019-01-09 | 2019-04-05 | 北京空间飞行器总体设计部 | A kind of optimal reconfiguration method under the J2 perturbation using Relative Navigation information |
| CN109613932A (en) * | 2019-01-09 | 2019-04-12 | 北京空间飞行器总体设计部 | Optimal continuous thrust control method under a kind of J2 perturbation using Relative Navigation information |
Non-Patent Citations (3)
| Title |
|---|
| BAICHUN GONG等: "Maneuver-free approach to range-only initial relative orbit determination for spacecraft proximity operations", 《ACTA ASTRONAUTICA》 * |
| 张朔等: "J2摄动下基于平均轨道要素差的集群航天器空间圆形编队设计", 《大连理工大学学报》 * |
| 龚柏春等: "空间非合作目标近程交会仅测角相对导航方法", 《中国惯性技术学报》 * |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113282096A (en) * | 2021-06-04 | 2021-08-20 | 中国人民解放军战略支援部队航天工程大学 | Control method for relative position nonlinear error of geostationary orbit game spacecraft |
| CN113282096B (en) * | 2021-06-04 | 2023-06-09 | 中国人民解放军战略支援部队航天工程大学 | Control method for nonlinear error of relative position of geostationary orbit game spacecraft |
| CN113761809A (en) * | 2021-11-08 | 2021-12-07 | 南京航空航天大学 | Passive detection orbit determination method based on deep neural network |
| CN113761809B (en) * | 2021-11-08 | 2022-04-08 | 南京航空航天大学 | Passive detection orbit determination method based on deep neural network |
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